1. Field of the Invention
The invention relates to aircraft gas turbine engines with counter rotating low pressure turbine rotors and, particularly, for such engines having inter-turbine frames that support the counter rotating low pressure turbine rotors in bearings and are used to mount the engines to the aircraft.
2. Description of Related Art
A gas turbine engine of the turbofan type generally includes a forward fan and booster compressor, a middle core engine, and an aft low pressure power turbine. The core engine includes a high pressure compressor, a combustor and a high pressure turbine in a serial flow relationship. The high pressure compressor and high pressure turbine of the core engine are interconnected by a high pressure shaft. The high pressure compressor, turbine, and shaft essentially form the high pressure rotor. The high pressure compressor is rotatably driven to compress air entering the core engine to a relatively high pressure. This high pressure air is then mixed with fuel in the combustor and ignited to form a high energy gas stream. The gas stream flows aft and passes through the high pressure turbine, rotatably driving it and the high pressure shaft which, in turn, rotatably drives the compressor.
The gas stream leaving the high pressure turbine is expanded through a second or low pressure turbine. The low pressure turbine rotatably drives the fan and booster compressor via a low pressure shaft, all of which form the low pressure rotor. The low pressure shaft extends through the high pressure rotor. Some low pressure turbines have been designed with counter rotating turbines that power counter rotating fans and booster or low pressure compressors. U.S. Pat. Nos. 4,860,537, 5,307,622 and 4,790,133 disclose counter rotating turbines that power counter rotating fans and booster or low pressure compressors. Most of the thrust produced is generated by the fan. Engine frames including fan and turbine frames are used to support and carry the bearings which, in turn, rotatably support the rotors. Bearing support frames are heavy and add weight, length, and cost to the engine.
Large modern commercial turbofan engines have higher operating efficiencies with higher bypass ratio configurations, larger transition ducts between low pressure and high pressure turbines. The frames, especially those located in the engine hot section, are complex and expensive. Other mid-size turbofan engines eliminate one frame by providing HP rotor support through a differential bearing arrangement in which the high pressure rotor rides on the low pressure rotor with an inter-shaft or differential bearing between them. New commercial engine designs are incorporating counter rotating rotors for improved turbine efficiency. Counter rotating rotors can have a detrimental impact on high pressure ratio component clearances, especially in the hot section which rely on tight clearance control to provide fuel efficiency benefits. Consequently, a need exists for engine and bearing support that will reduce engine length, weight, and cost and still provide low tip clearance losses.
An aircraft gas turbine engine assembly includes a high pressure rotor including a high pressure turbine, a low pressure turbine having counter rotating low pressure inner and outer rotors located aft of the high pressure rotor, and an inter-turbine frame axially located between the high and low pressure turbines. The low pressure inner and outer rotors including low pressure inner and outer shafts which are at least in part rotatably disposed co-axially with and radially inwardly of the high pressure rotor. The inter-turbine frame has a first structural ring, a second structural ring disposed co-axially with and radially spaced inwardly of the first structural ring about a centerline, and a plurality of circumferentially spaced apart struts extending radially between the first and second structural rings. Forward and aft sump members, having forward and aft central bores respectively, are fixedly joined to axially spaced apart forward and aft portions of the inter-turbine frame by forward and aft bearing support structures, respectively. The low pressure inner and outer rotors are rotatably supported by an aftwardmost low pressure rotor support bearing mounted in the aft central bore of the aft sump member. The high pressure rotor is aftwardly radially rotatably supported by a fifth bearing mounted in the forward bearing support structure. A frame connecting means for connecting the engine to an aircraft is located on the first structural ring. In an exemplary embodiment of the invention, the connecting means includes at least one U-shaped clevis.
An outlet guide vane assembly supports a row of outlet guide vanes that extend radially between a low pressure turbine casing structurally connected to the inter-turbine frame and an annular box structure. A cover plate is bolted to the annular box structure. A rotatable annular outer drum rotor is drivingly connected to a first fan blade row and a first booster by the low pressure inner shaft. A rotatable annular inner drum rotor is drivingly connected to a second fan blade row and a second booster by the low pressure outer shaft, the first and second boosters are axially located between the first and second fan blade rows.
A bypass duct radially bounded by a fan casing and an annular radially inner bypass duct wall surrounds the first and second boosters and a radially outer portion of the second fan blade row is radially disposed within the bypass duct. The engine assembly has a fan inlet hub to tip radius ratio in a range between 0.20 and 0.35, a bypass ratio in a range of 5-15, an operational fan pressure ratio in a range of 1.4-2.5, and a sum of operational fan tip speeds of the first and second fan blade rows in a range of 1000 to 2500 feet per second. The high pressure compressor is designed and operable to produce a compressor pressure ratio in a range of about 15-30 and overall pressure ratio in a range of about 40-65.
Further embodiments of the invention include a second seal in sealing arrangement between forward ends of the low pressure turbine casing and the outer drum rotor, a third seal in sealing arrangement between the low pressure turbine casing and a final stage of the low pressure turbine blade rows which is bolted to an aft end of the outer drum rotor, and a first seal in sealing arrangement between the second fan and the fan frame. The seals are brush seals, however in other embodiments the seals may be non-contacting seals or a combination of brush seals and non-contacting seals. The non-contacting seals may be aspirating seals or face seals.